In flight, an aircraft produces pressure waves or disturbances in the air through which it is flying. These pressure waves propagate at the speed of sound. When the aircraft flies at subsonic speed, these pressure waves propagate in all directions around the aircraft, including ahead of the aircraft. When aircraft fly at supersonic speed, these pressure waves cannot propagate ahead of the aircraft because the aircraft is traveling faster than the propagation speed of the waves. Instead, the pressure waves generated by the aircraft typically coalesce into two shock waves, one associated with the nose of the aircraft and the other associated with the tail of the aircraft. These shock waves pressure differentials that propagate circumferentially away from the aircraft. With respect to the shock wave associated with the nose (the “bow shock”), the pressure increases abruptly from about ambient to a pressure significantly thereabove. The pressure decreases down from this pressure significantly above ambient down to a pressure below ambient in the region between the bow shock and the shock wave associated with the tail (the “tail shock”). The pressure then increases abruptly from below ambient to about ambient at the tail shock.
These shock waves can propagate great distances away from the aircraft and eventually those that are directed downwardly will reach the ground where they can produce significant acoustic disturbances called sonic booms. Sonic booms are so named because of the sounds created by the abrupt pressure changes when the shock waves pass a reference point on the ground. The acoustic signature of a sonic boom is commonly characterized as an N-wave because the pressure changes associated with the acoustic signature resemble the letter “N” when plotted as a function of position from the nose of the aircraft. That is, an N-wave is characterized by the abrupt pressure rise associated with the bow shock, commonly referred to as “peak overpressure,” followed by a decrease to a pressure below ambient. This is followed by the abrupt rise back toward ambient pressure in association with the tail shock. Where perceivable, typically on the ground by a person, a sonic boom effect is caused by the two rapid-succession, high magnitude pressure changes. Strong sonic booms cause an objectionably loud noise, as well as other undesirable conditions at ground level. For these reasons, supersonic flight over some populated areas is restricted. A schematic representation of the phenomenon of aircraft produced sonic boom is provided in FIG. 20.
It should be appreciated that shock waves propagate in the form of a “Mach Cone” having a shape defined by a Mach angle (μ). The Mach angle μ is a function of the Mach number M, which is defined as the ratio of the speed of an object over the speed of sound. The Mach angle (μ) can be determined using the equation:sin(μ)=1/M, orμ=sin−1(1/M)
The shape of the Mach cone produced by an aircraft in supersonic flight can be represented by rotating a line drawn from the aircraft's nose tip toward the tail of the aircraft and oriented at an angle (μ) with respect to the aircraft's direction of travel. Consequently, the tip of the Mach cone points in the direction of travel.
In order for supersonic flight over land to be acceptable, the pressure disturbances that cause the sonic boom's acoustic signature must be controlled to avoid strong sonic boom effects caused by the abrupt pressure changes due to the bow and tail shock waves.
It should be appreciated that it is not only the magnitude of the created pressures that are radiated to ground level from an aircraft flying at supersonic speeds that causes persons to experience unpleasant sonic boom effects, but it is primarily the rate(s) of change in the pressures experienced at ground level (pressure differentials−Δ P) that produces the undesirable sonic boom effects. Therefore, one goal for minimizing audible sonic boom effects is to control pressure differentials caused at ground level by a supersonic flying craft.
Another characteristic of the pressure waves or disturbances generated by a supersonic flying aircraft is that the elevated pressures associated essentially with the forward portion of the craft have an effect that coalesces together as they travel toward the ground. As FIG. 20 depicts, the lowered pressures associated essentially with the rearward portion of the craft also have an effect that coalesces together as they travel toward the ground. As described above, it is these two primary pressure changes that cause the sonic boom effects at ground level. Therefore, it can be a solution to the sonic boom problem to smooth the pressure differentials so that there are no abrupt changes. That is to say, the magnitude of the different pressures induced by a supersonic flying aircraft need not necessarily be altered, but it can be enough for some aircraft designs to smooth the abrupt pressure changes experienced at ground level to be more gradual.
Features of the aircraft that cause such abrupt changes in the induced pressures are also detrimental. As explained hereinabove, the pressure disturbances or waves radiate from the aircraft at a relationship based at least in part on the speed of the craft. The angle of radiation can also be affected by the magnitude of the caused disturbance. That is to say, and is best illustrated in FIG. 21, abrupt projections off of the fuselage of the aircraft (transverse to the direction of travel of the aircraft) will cause larger and higher angle disturbances than smooth transitions. In the case of FIG. 21, the outwardly projecting jet engines cause pressure waves; one at the top, forward projecting portion of the inlet, and another at the lower lip of the engine's inlet. The pressure disturbances induced by the engine of the aircraft in FIG. 21 coalesce and thereby detrimentally create a combined pressure differential at the ground. Therefore, working toward the goal of minimizing differentials in the pressure profile or signature of a supersonic aircraft, a design challenge has been identified to keep transverse projections (to the direction of travel of the aircraft), and even surface disruptions to a minimum. In this context, a surface disruption is considered to be any dimensional change along the length of the aircraft that is transverse to the axis of travel. Since it is pressure waves radiating from the bottom of the plane that most effects ground boom, it is to the extreme lower surfaces of the aircraft that this smoothing goal is most relevant.
As background to the present invention(s), it is known that attempts have been made to modify the design of supersonic aircraft in order to adjust the sonic boom signature. These modifications have included changes to wing design, as described in U.S. Pat. No. 5,934,607, issued to Rising, et al., for a “Shock Suppression Supersonic Aircraft.” Another approach involves incorporating air passages through the fuselage or wings of supersonic aircraft, such as the structures described in U.S. Pat. No. 4,114,836, issued to Graham, et al., for an “Airplane Configuration Design for the Simultaneous Reduction of Drag and Sonic Boom”; U.S. Pat. No. 3,794,274, issued to Eknes, for an “Aircraft Structure to Reduce Sonic Boom Intensity”; and U.S. Pat. No. 3,776,489, issued to Wen, et al., for a “Sonic Boom Eliminator.” Further attempts at reducing the sonic boom caused by supersonic aircraft include the addition to the aircraft of structure arranged to disrupt the air flow patterns as the aircraft travels at supersonic speed. Examples include the structure described in U.S. Pat. No. 3,709,446, issued to Espy, for a “Sonic Boom Reduction” and U.S. Pat. No. 3,647,160, issued to Alperin, for a “Method and Apparatus for Reducing Sonic Booms.”
Another attempt to control the sonic boom in a supersonic aircraft uses a blunt nose to increase the air pressure immediately adjacent to the nose of the aircraft, thus disrupting the normal formation of the pressure wave that causes the acoustic signature. This disruption results in a reduction of the abruptness of the pressure changes that develop after the initial pressure rise in the acoustic wave that strikes the ground. A blunt nose, however, does not reduce the initial overpressure rise in the resulting boom signature. Furthermore, a blunt nose creates a significant amount of drag on the aircraft, drastically decreasing its efficiency.
U.S. Pat. No. 5,740,984, issued to Morgenstern, for a “Low Sonic Boom Shock Control/Alleviation Surfaces” describes a mechanical device on the nose of the airplane which can be moved between a first position effecting a blunt nose when sonic boom reduction is desired and a second position effecting a streamlined nose when sonic boom reduction is not required, thereby removing (in the streamlined configuration) the drag penalty inherent in a blunt nose design.
U.S. Pat. Nos. 5,358,156, 5,676,333, and 5,251,846, all issued to Rethorst and all entitled “Supersonic Aircraft Shock Wave Energy Recovery System” (collectively “the Rethorst patents”), describe an aircraft with a modified wing design and a forward ring on the fuselage for eliminating the sonic boom of a supersonic aircraft. FIG. 19 in each of the Rethorst patents shows a side elevation view of an aircraft whose nose coincides with the bottom of its fuselage. It appears from FIGS. 19A and 19B that the bottom of at least a portion of the fuselage is planar. The Rethorst patents do not provide further disclosure regarding this fuselage shape, and they do not teach non-uniform propagation of pressure disturbances about the fuselage. To the contrary, the Rethorst patents teach that the initial bow shock is axisymmetric about the nose. See U.S. Pat. No. 5,676,333 at col. 14, lines 31-34; U.S. Pat. No. 5,738,156 at col. 14, lines 6-10; and U.S. Pat. No. 5,251,846 at col. 14, lines 9-12.
Regarding another aspect of the present invention, the same being the inclusion of a leading and/or trailing spike on the supersonic aircraft, the Rethorst patents also describe a supersonic aircraft having a spike extending from the front of the aircraft and a forward ring on the fuselage for eliminating a sonic boom. The spike is described to direct the bow shock onto the manifold ring that recovers the shock energy and converts it to useful work. The spike is further depicted as being extendable, but it does not include a complex surface contour, and it is not disclosed to include a number of (plurality) telescopically collapsible sections. Instead, the Rethorst spike is disclosed as being a single cylindrical member that tapers to a point at a leading end.
U.S. Pat. No. 4,650,139, issued to Taylor et al., discloses a blunt-nosed spike that can be extended from a space vehicle's fuselage.
U.S. Pat. No. 3,643,901, issued to Patapis, discloses a ducted spike for attachment to a blunt body operating at supersonic speed for the purpose of receiving and diffusing oncoming air to reduce pressure drag on, and erosion of the blunt body.
U.S. Pat. No. 3,425,650, issued to Silva, discloses an apparatus that can be extended on a boom from the front of an aircraft to deflect air outwardly therefrom.
U.S. Pat. No. 3,655,147, issued to Preuss, covers a device attached to the lower forebody of an aircraft for the purpose of reflecting pressure disturbances caused by the aircraft's flight in directions away from the ground.
Although some of the foregoing documents are directed to sonic boom mitigation, none of them address the sonic boom signature shaping techniques of the present invention.